The invention is more particularly applicable for use in spatial telecommunications applications.
Although the invention is applicable to any type of satellite and any type of orbit, it shall subsequently be considered for satellites with an elliptical orbit inclined with respect to the equatorial plane.
By way of example, this shall thus relate to satellites intended for a communication system denoted by the Applicant by the term "SYCOMORES", this abbreviation standing for a "System of Communication with Mobile Units Relayed by Satellites".
Such a communication system is described in the French patent No 88/02632 filed by the Applicant on 2 Mar. 1988.
FIG. 1 briefly summarizes the essential elements of this system.
A central station SC is situated at the center of a spherical triangle-shaped covering zone Z. Two geosynchronous satellites S-A and S-B have elliptical orbits with identical parameters. By way of explanation, these parameters may be the following:
apogee situated at about 50,543.4 km, PA0 perigee situated at about 21,028.6 km, PA0 meniscal axis of 42,164 km, PA0 inclination of 63.degree. 4, PA0 perigee argument 270, PA0 orbit excentricity 0.35.
Each satellite includes an antenna or antennae 11 and 12; each antenna is orientated towards the central station throughout the period when the satellite moves above the covering zone.
The central station includes one connection station and one control station.
FIG. 1 also shows a mobile unit M (which of course is situated inside the zone Z but which is shown above the latter for the sake of more clarity). This mobile unit is equipped with an antenna 14 whose axis continuously points towards the zenith.
In order to station such satellites, a large number of strategies are possible, but only one shall be described with reference to FIG. 2. This stationing uses the ARIANE IV rocket and requires three pulses. At the time of launching, the satellite is accompanied by an ordinary geostationary satellite. The two satellites are placed on the standard transfer orbit of the ARIANE IV rocket, this orbit being situated within a quasi-equatorial plane (inclination of 7.degree.) with a perigee at 200 km, an apogee at 35,975 km and a perigee argument of 178.degree. (orbit marked OST on FIG. 2).
Close to the perigee, a satellite rocket is ignited for a first pulse suitable for raising the apogee to 98,000 km, the orbit remaining within the same plane (orbit 01). This pulse may be broken down into two or three pulses.
Close to the apogee of the orbit 01, a new pulse is sent to the satellite to change the plane of its orbit. The inclination of this plane is close to that of the plane of the definitive orbit (namely 63.degree. 4). This thrust is the largest and may be broken down into two or three thrusts. The orbit then becomes 02.
Finally, at an appropriate point of this orbit, a third thrust is sent to the satellite so as to provide it with a definitive orbit.
If this strategy is satisfactory in certain respects, it nevertheless does constitute a drawback. In fact, it requires that the orbital plane be tilted when passing from the orbit 01 to the orbit 02, this resulting in a considerable consumption of propellant. It would be preferable to be able to reduce this consumption so as to increase the mass of the carrying capacity.